Field of the Invention
Present embodiments relate generally to gas turbine engines. More particularly, but not by way of limitation, present embodiments relate to ceramic matrix composite combustor liners.
Description of the Related Art
A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet or intake is at a forward end of the engine. Moving toward the aft end of the engine, in order, the intake is followed in serial flow communication by a compressor, a combustion chamber and a turbine. It will be readily apparent from those skilled in the art that additional components may also be included in the gas turbine engine, such as, for example, low-pressure and high-pressure compressors, and high-pressure and low-pressure turbines. This, however, is not an exhaustive list. An engine also typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The internal shaft is connected to both the turbine and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades.
In operation, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages utilize blades to extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. In a multi-stage turbine, a second stage stator nozzle assembly is positioned downstream of the first rotor stage blades followed in turn by a row of second stage turbine rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy and drive the shaft turning the high pressure compressor. One or more stages of a low pressure turbine may be mechanically coupled to a low pressure or booster compressor for driving the booster compressor and additionally an inlet fan.
In driving improvement of engine operating efficiency, it has been a desired goal to increase operating temperatures within the engine. However, one obstacle has been material temperature limitation which must be kept below critical levels. Otherwise, the material or component formed from the material may be damaged. One promising material has been ceramic matrix composite due to its lightweight, formability and ability to operate at extremely high temperatures associated with turbine engines. For example, in the area of combustor development, the combustor must be capable of meeting the design life requirements for use in the turbine engine operating temperature environment. The use of ceramic matrix composite (CMC) is desirable due to its temperature resistance characteristics. To enable combustor liners to operate effectively in such strenuous temperature conditions, it has been practiced to utilize composite and, in particular, ceramic matrix composite (CMC) materials for use in the shroud segments because they have higher temperature capability than metallic type parts. However, such ceramic matrix composites (CMC) have mechanical characteristics that must be considered during the design and application of the CMC combustor liners. CMC materials have a coefficient of thermal expansion which differs significantly from metal alloys used to form the combustor and to which the combustor liner is connected. Therefore, if a CMC component is restrained and cooled on one surface during operation, stress concentrations can develop leading to failure of the component. Additionally, vibration can lead to wear as well as problems with leakage about the combustor liner, all of which result in inefficient operation of the combustor.
As may be seen by the foregoing, it would be desirable to allow the use of ceramic matrix composites within the combustor so as to allow higher operating temperatures and more efficient gas turbine engine operation while compensating for the above operating condition and criteria.
The information included in this Background section of the specification, including any references cited herein and any description or discussion thereof, is included for technical reference purposes only and is not to be regarded subject matter by which the scope of the invention is to be bound.